EP0273849B1 - Inlet total temperature synthesis for gas turbine engines - Google Patents

Inlet total temperature synthesis for gas turbine engines Download PDF

Info

Publication number
EP0273849B1
EP0273849B1 EP87630243A EP87630243A EP0273849B1 EP 0273849 B1 EP0273849 B1 EP 0273849B1 EP 87630243 A EP87630243 A EP 87630243A EP 87630243 A EP87630243 A EP 87630243A EP 0273849 B1 EP0273849 B1 EP 0273849B1
Authority
EP
European Patent Office
Prior art keywords
engine
signal
temperature
indicative
signal indicative
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired
Application number
EP87630243A
Other languages
German (de)
English (en)
French (fr)
Other versions
EP0273849A3 (en
EP0273849A2 (en
Inventor
Leon Krukoski
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
RTX Corp
Original Assignee
United Technologies Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP0273849A2 publication Critical patent/EP0273849A2/en
Publication of EP0273849A3 publication Critical patent/EP0273849A3/en
Application granted granted Critical
Publication of EP0273849B1 publication Critical patent/EP0273849B1/en
Expired legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/26—Control of fuel supply
    • F02C9/28—Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2270/00—Control
    • F05D2270/01—Purpose of the control system
    • F05D2270/11—Purpose of the control system to prolong engine life
    • F05D2270/112—Purpose of the control system to prolong engine life by limiting temperatures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2270/00—Control
    • F05D2270/30—Control parameters, e.g. input parameters
    • F05D2270/303—Temperature

Definitions

  • This invention relates to controls for gas turbine engines and more particularly to synthesizing parameters for use in such controls.
  • the most current gas turbine engines utilize electronic controls which automatically regulate engine operation, such as fuel flow rate and compressor bleed position, based upon pilot demand (i.e., throttle position), various aircraft parameters (e.g., aircraft speed and altitude), engine parameters (e.g., burner pressure and exhaust gas temperature), and empirically developed relationships between various parameters.
  • the empirically developed relationships are built into the control in the form of schedules. Considerable redundancy is built into the control system to maintain engine operation as close to normal as possible and to prevent unnecessary engine shutdowns in the event that certain parameters upon which the control primarily depends cannot be accurately determined or calculated, such as due to faulty measuring equipment, circuit malfunction, or for any other reason.
  • burner pressure is synthesized by generating a ratio of burner pressure to total engine inlet pressure from known values of total engine inlet temperature, compressor speed, and an empirically developed relationship therebetween for that particular engine. Similarly, total inlet pressure is estimated from known relationships between aircraft altitude and Mach number. Multiplying the calculated total inlet pressure times the ratio of burner pressure to total inlet pressure yields a synthesized value of the burner pressure.
  • An important parameter for engine controls is the total inlet temperature to the engine. This is an aircraft parameter as opposed to an engine parameter, such as burner pressure, since inlet temperature is essentially unaffected by engine operation. Generally total inlet temperature is measured at one or more locations at the engine inlet by means of temperature probes. These probes are vulnerable to bird strikes which can damage them. Additionally, ice accumulation within the probes can cause them to generate faulty readings. Until now, if no good reading for the engine inlet temperature can be obtained, the control reverts to an alternate control mode which performs on the basis of the last reliable inlet temperature reading. This may eventually require the engine to be shut down even though the engine may be capable of operating in a perfectly normal fashion.
  • One object of the present invention is a gas turbine engine control system which synthesizes the value of an aircraft (as opposed to engine) parameter.
  • Another object of the present invention is a gas turbine engine control system which generates a synthesized value of total engine inlet temperature.
  • Another object of the present invention is a gas turbine engine control system which uses a synthesized engine total inlet temperature as a control parameter when the actual inlet temperature measurement is unreliable.
  • a gas turbine engine control system generates, as a function of aircraft Mach number, a signal which is indicative of a ratio of a measured engine parameter and the engine total inlet temperature, and combines that ratio with a signal indicative of the measured engine parameter to generate an output signal indicative of the approximate value of the engine total inlet temperature.
  • the approximate or synthesized engine total inlet temperature is utilized as the control parameter in place of the measured temperature.
  • the ratio of the engine exhaust nozzle pressure (P s ) to ambient pressure (Pamb) is used, in conjunction with the aircraft Mach number, in a function generator to produce an empirically developed ratio indicative of the temperature rise from the inlet to the outlet of the compressor (Ta/T z ). (In a twin spool engine this would be the temperature rise ratio across both the high and low compressor.) This ratio may then be multiplied or divided, as appropriate, by the measured temperature at the compressor outlet (T 3 ) to yield a synthesized value of the temperture at the compressor inlet (T 2 syn).
  • Mach number is approximated or synthesized from a schedule based upon aircraft altitude.
  • the empirically derived relationship between the temperature rise ratio Ta/T 2 , Mn and the nozzle expansion ratio Ps/Pamb used in the synthesis of the engine inlet temperature is not sufficiently accurate. Since engine inlet temperature T 2 is a relatively slow changing parameter, during transient engine modes of operation the control uses the synthesized value of the engine inlet temperature calculated just prior to the engine entering the transient mode.
  • the sole figure is a schematic and block diagram of a twin spool gas turbine engine incorporating the control system of the present invention.
  • twin spool gas turbine engine shown in the drawing and generally represented by the reference numeral 10.
  • the engine comprises a low compressor 12 connected through a shaft to a low turbine 14; a high compressor 16 connected through a shaft to a high turbine 18; and a burner section 20 disposed between the high compressor and high turbine.
  • An electronic engine control automatically regulates engine operation, such as fuel flow rate and compressor bleed position, based upon pilot demand, various aircraft and engine parameters, and scientific and empirically developed relationships between various parameters.
  • the electronic engine control uses the total engine inlet temperature as a key parameter for automatically controlling the engine.
  • One aspect of the electronic engine control is to generate a synthesized value of the total engine inlet temperature for use in the event of the unavailability or unreliability of an actually measured inlet temperature. It is that portion of the electronic engine control which is shown in the drawing. Referring to the drawing, a signal 22 indicative of the airplane Mach number is delivered to a switch 24. Mach number is typically determined by measuring the compressor inlet pressure by means of a probe and then calculating the Mach number through a known scientific relationship between Mach number and pressure. A signal 26 is delivered to the switch if it is determined, by means not shown, that the pressure measurement used to generate the Mach number is available and reliable. In that case the signal 22 is passed through the switch 24 to a function generator 28.
  • a synthesized value of the Mach number is generated by a function generator 30 as a function of ambient pressure Pamb, which is measured and delivered to the function generator via the line 32.
  • a signal 34 indicative of this synthesized Mach number is delivered from the function generator 30 into the switch 24 and is passed through to the function generator 28 when no signal 26 is being sent to the switch indicating that the Mach number signal 22 is either unavailable or unreliable.
  • the engine exhaust pressure P 5 is measured and a signal 36 indicative thereof is provided to a divider 38 along with a signal 39 indicative of P a mb.
  • the divider generates a signal 40 which is the ratio of P 5 to P amb , which ratio is generally referred to as the nozzle expansion ratio.
  • the function generator 28 generates a signal 42 indicative of the ratio of the temperature rise across the compressors 12, 16. That ratio is herein designated as T 3 /T 2 . That signal is ultimately delivered to a divider 46 as the numerator thereof.
  • the actual high compressor outlet temperature T 3 is measured and a signal 44 indicative thereof is also sent to a divider 46 as the denominator thereof. Because of the time it takes for a temperature probe to respond to temperature changes, the value of T 3 delivered into the divider 46 will be the temperature which existed shortly before that temperature value was delivered to the divider 46. Therefore, the temperature rise ratio signal 42 is passed through a device 48 which has a built-in time delay such that the temperature rise ratio signal 50 delivered therefrom into the divider 46 is the temperature rise calculated at the time the temperature T 3 delivered into the divider 46 was actually measured. Such time delays are well known in the art for many applications in control systems.
  • the divider 46 delivers a signal 52 indicative of the inverse of the approximate total engine inlet temperature.
  • This signal is passed through a calculator 54 which calculates the inverse of the value of the signal 52, thereby producing a signal 56 indicative of the approximate or synthesized total engine inlet temperature T 2s y n .
  • the signal 56 is delivered to a switch 58.
  • the temperature rise ratio 42 generated by the function generator 28 calculated during such transient engine operations are not used.
  • the high rotor speed derivative N 2 is calculated or otherwise determined, and its absolute value is delivered into a threshhold detector 60. If the high compressor speed is changing at a rate greater than a predetermined value x, then a positive or true signal 62 is delivered to the switch 58. Otherwise no signal is sent to the switch.
  • the output of switch 58 is a signal 66 which is indicative of the value of the synthesized temperature signal 56 last calculated prior to the signal 62 being positive (i.e., before the engine went into the transient mode of operation). And that signal 56 is maintained constant until the signal 62 ceases. When no signal 62 is present the switch 58 simply passes the presently calculated value of signal 56.
  • the signal 68 is indicative of the synthesized value of the inlet temperature passed by the switch 58, and is delivered into another switch 70 along with a signal 72 indicative of the actual total engine inlet temperature T 2 as measured by a probe at the inlet to the low compressor 12. Via means not shown it is determined whether the measured temperature signal 72 is either unavailable or unreliable. If the answer is "yes”, a positive or true signal 74 is delivered to the switch 70, and the switch output signal 76 will be the synthesized total inlet temperature signal 68. Otherwise, the switch output signal 76 will be the acual mesaured temperature signal 72. In either case, the signal 76, whether indicative of the actual or synthesized inlet temperature, is the signal used as a control parameter in the elecronic engine control.
  • the temperature rise ration Ta/T 2 is not the only temperature ratio which may be used in the present invention. Any temperature ration which is a ratio of the temperature T 2 and a temperature within the engine and which can be calculated from empirical relationships between it, the aircraft Mach number, and other determinable parameters, such as nozzle expansion ratio or the like, may be used. That temperature ration would thereupon be either divided or multiplied by the appropriate measured temperature to yield either the synthesized turbine inlet temperature or its inverse, as the case may be.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Control Of Turbines (AREA)
  • Electrical Control Of Air Or Fuel Supplied To Internal-Combustion Engine (AREA)
EP87630243A 1986-12-08 1987-11-19 Inlet total temperature synthesis for gas turbine engines Expired EP0273849B1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US06/939,218 US4748804A (en) 1986-12-08 1986-12-08 Inlet total temperature synthesis for gas turbine engines
US939218 1986-12-08

Publications (3)

Publication Number Publication Date
EP0273849A2 EP0273849A2 (en) 1988-07-06
EP0273849A3 EP0273849A3 (en) 1988-07-20
EP0273849B1 true EP0273849B1 (en) 1990-07-11

Family

ID=25472764

Family Applications (1)

Application Number Title Priority Date Filing Date
EP87630243A Expired EP0273849B1 (en) 1986-12-08 1987-11-19 Inlet total temperature synthesis for gas turbine engines

Country Status (5)

Country Link
US (1) US4748804A (ja)
EP (1) EP0273849B1 (ja)
JP (1) JP2644785B2 (ja)
CA (1) CA1286774C (ja)
DE (1) DE3763655D1 (ja)

Families Citing this family (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5379584A (en) * 1992-08-18 1995-01-10 Alliedsignal Inc. Synthesis of critical temperature of a turbine engine
US5394689A (en) * 1993-09-22 1995-03-07 General Electric Company Gas turbine engine control system having integral flight Mach number synthesis method
GB9410760D0 (en) * 1994-05-27 1994-07-27 Rolls Royce Plc Gas turbine engine fuel control system
US5622042A (en) * 1995-02-27 1997-04-22 Compressor Controls Corporation Method for predicting and using the exhaust gas temperatures for control of two and three shaft gas turbines
US5775090A (en) * 1996-12-23 1998-07-07 Allison Engine Company Torque signal synthesis method and system for a gas turbine engine
US5775089A (en) * 1996-12-23 1998-07-07 Allison Engine Company Pressure signal synthesis method and system for a gas turbine engine
JP4146049B2 (ja) 1999-10-05 2008-09-03 本田技研工業株式会社 航空機用ガスタービン・エンジンの制御装置
JP2002309963A (ja) * 2001-04-17 2002-10-23 Mitsubishi Heavy Ind Ltd ガスタービンプラント
JP4529521B2 (ja) * 2004-04-05 2010-08-25 株式会社Ｉｈｉ 圧縮機用翼揺動制御装置、ファン用翼揺動制御装置、圧縮機、及びファン
DE102005017400B4 (de) * 2005-04-15 2012-03-15 Mtu Aero Engines Gmbh Verfahren zur Bestimmung der Turbineneintrittstemperatur einer Gasturbine bei instationären Vorgängen
US20100003123A1 (en) * 2008-07-01 2010-01-07 Smith Craig F Inlet air heating system for a gas turbine engine
US8915088B2 (en) * 2010-06-11 2014-12-23 Hamilton Sundstrand Corporation Fuel control method for starting a gas turbine engine
US8560203B2 (en) * 2010-07-30 2013-10-15 Pratt & Whitney Canada Corp. Aircraft engine control during icing of temperature probe
US9828106B2 (en) * 2015-06-18 2017-11-28 Honeywell International Inc. Aircraft gas turbine propulsion engine control without aircraft total air temperature sensors
CN114013684B (zh) * 2021-11-15 2024-05-07 中国航发沈阳发动机研究所 一种新研航空发动机持久试车进气温度确定方法及装置

Family Cites Families (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2809492A (en) * 1952-12-23 1957-10-15 Simmonds Aerocessories Inc Apparatus for measuring and/or controlling fuel/air ratio of gas turbines without direct gravimetric fuel metering
GB1135614A (en) * 1966-02-23 1968-12-04 Rolls Royce Fuel control system for a gas turbine engine
US3377848A (en) * 1966-08-22 1968-04-16 Gen Electric Temperature indicating means for gas turbine engines
US3789665A (en) * 1972-02-22 1974-02-05 Avco Corp Inferred measurement of the turbine inlet temperature of a gas turbine engine
JPS5722027Y2 (ja) * 1975-12-08 1982-05-13
US4058975A (en) * 1975-12-08 1977-11-22 General Electric Company Gas turbine temperature sensor validation apparatus and method
GB2011091B (en) * 1977-12-22 1982-04-28 Gen Electric Method and apparatus for calculating turbine inlet temperature
US4228650A (en) * 1978-05-01 1980-10-21 United Technologies Corporation Simulated parameter control for gas turbine engine
US4212161A (en) * 1978-05-01 1980-07-15 United Technologies Corporation Simulated parameter control for gas turbine engine
US4543782A (en) * 1982-05-21 1985-10-01 Lucas Industries Gas turbine engine fuel control systems
US4581888A (en) * 1983-12-27 1986-04-15 United Technologies Corporation Compressor rotating stall detection and warning system
US4651518A (en) * 1984-12-18 1987-03-24 United Technologies Corporation Transient derivative scheduling control system
US4594849A (en) * 1984-12-20 1986-06-17 United Technologies Corporation Apparatus for synthesizing control parameters

Also Published As

Publication number Publication date
JPS63150436A (ja) 1988-06-23
DE3763655D1 (de) 1990-08-16
CA1286774C (en) 1991-07-23
JP2644785B2 (ja) 1997-08-25
EP0273849A3 (en) 1988-07-20
US4748804A (en) 1988-06-07
EP0273849A2 (en) 1988-07-06

Similar Documents

Publication Publication Date Title
EP0273849B1 (en) Inlet total temperature synthesis for gas turbine engines
US4060980A (en) Stall detector for a gas turbine engine
US4117668A (en) Stall detector for gas turbine engine
US4622808A (en) Surge/stall cessation detection system
EP0005135B1 (en) Exhaust nozzle control and core engine fuel control for turbofan engine
US3935558A (en) Surge detector for turbine engines
US4767259A (en) Cooling air flow controlling apparatus for gas turbine
EP0187115B1 (en) Apparatus for synthesizing control parameters
US4060979A (en) Stall warning detector for gas turbine engine
US4581888A (en) Compressor rotating stall detection and warning system
US5197280A (en) Control system and method for controlling a gas turbine engine
EP0187114B1 (en) Transient gas turbine engine bleed control
EP0324633A2 (en) Fuel control system
US4212161A (en) Simulated parameter control for gas turbine engine
EP1069296B1 (en) A method of obtaining an indication of the power output of a turbine
KR20000047986A (ko) 가스 터빈 엔진 제어 시스템
US4894782A (en) Diagnostic system for determining engine start bleed strap failure
US6915639B1 (en) Method and apparatus for gas turbine over-speed protection
US4248042A (en) Engine thrust control system
US5072580A (en) System for operating gas turbine jet engine with fan damage
EP0401152B1 (en) Acceleration control for a gas turbine engine with duct pressure loss compensation
EP0385912B1 (en) Gas turbine jet engine
RU2319025C1 (ru) Способ управления газотурбинным двигателем
EP2975479B1 (en) Method for the control and protection of a gas turbine and gas turbine using such method
GB2229556A (en) Controlling a gas turbine engine

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

PUAL Search report despatched

Free format text: ORIGINAL CODE: 0009013

AK Designated contracting states

Kind code of ref document: A2

Designated state(s): DE FR GB IT

AK Designated contracting states

Kind code of ref document: A3

Designated state(s): DE FR GB IT

17P Request for examination filed

Effective date: 19881228

17Q First examination report despatched

Effective date: 19890602

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): DE FR GB IT

ET Fr: translation filed
REF Corresponds to:

Ref document number: 3763655

Country of ref document: DE

Date of ref document: 19900816

ITF It: translation for a ep patent filed
PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

26N No opposition filed
ITTA It: last paid annual fee
PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: FR

Payment date: 19931011

Year of fee payment: 7

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: GB

Payment date: 19931013

Year of fee payment: 7

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: DE

Payment date: 19931020

Year of fee payment: 7

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: GB

Effective date: 19941119

GBPC Gb: european patent ceased through non-payment of renewal fee

Effective date: 19941119

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: FR

Effective date: 19950731

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: DE

Effective date: 19950801

REG Reference to a national code

Ref country code: FR

Ref legal event code: ST

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: IT

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES;WARNING: LAPSES OF ITALIAN PATENTS WITH EFFECTIVE DATE BEFORE 2007 MAY HAVE OCCURRED AT ANY TIME BEFORE 2007. THE CORRECT EFFECTIVE DATE MAY BE DIFFERENT FROM THE ONE RECORDED.

Effective date: 20051119